Fixed gain passive adaptive aircraft control system

ABSTRACT

AN AIRCRAFT CONTROL SYSTEM INCLUDING A CONSTANT HIGH-GAIN FEEDBACK LOOP FOR FORCING AIRCRAFT RESPONSE TO FOLLOW A DESIRED RESPONSE AND THEREBY INSURING ADEQUATE HANDLING STABILITY OVER A WIDE RANGE OF FLIGHT CONDITIONS. ACCELERATION SENSORS ARE LOCATED FORWARD OF THE CENTER OF GRAVITY OF THE CRAFT FOR SENSING ANGULAR AS WELL AS LINEAR ACCELERATION TO PERMIT A HIGHER GAIN THEN WOULD OTHERWISE BE THE CASE AND A FILTER PROVIDES A HIGHER GAIN IN THE FREQUENCY RANGE OF THE DESIRED AIRCRAFT RESPONSE TO IMPROVE SYSTEM PERFORMANCE WITHOUT APPRECIABLY COMPROMISING STABILITY IN OTHER FREQUENCY RANGES.

United States Patent [72]. Inventor Erwin A. Nlumann 3,0575 84 10/1962Bretoi 318120050X Lodl, NJ. 3,345,018 10/1967 Chanak et al.... 244/77(F)[21] Appl. No. 787,889 3,399,849 9/1968 Hendrick 244/77(D) [22] 1968Primar ExaminerMilton Buchler 45 P 1 ted June 28 1971 Y l 3 E k C onAssistant Examiner-Jeffrey L. Forman [73 The and AttorneysAnthony FCuoco and Plante, Hartz, Smith 8!.

Thompson [54] FIXED GAIN PASSIVE ADAPTIVE AIRCRAFT CONTROL SYSTEM 11 Chi2 Drlwln F m I b- ABS'IRACT: An aircraft control system including aconstant 244/77, high-gain feedback loop for forcing aircraft responseto follow 318/18 a desired response and thereby insuringadequatehandling [51] Int. Cl. B64c 13/18 stability over a wide range f fli ht ii Acceleration of Selrch sensors are located forward of the center ofgravity of the craft E- -Z -4 18 for sensing angular as well as linearacceleration to permit a 1 higher gain then would otherwise be the caseand a filter pro- [56] References Chad vides a higher gain in thefrequency range of the desired air- UNITED STATES PATENTS craft responseto improve system performance without ap- 3,051,416 8/1962 Roticr244/77(G) preciably compromising stability in other frequency ranges.

LINKAGE I i l I 2 5 24 F I I I; I as I 6 3o 1 FORCE PITCH SURFACE 4-DAMP R v&

senson SEW?) "'ACI'UATOR MR CRAFT I 14 I PITCH RATE GYRO J l I 22 NORMALACCELEROMETER FIXED GAIN PASSIVE ADAPTIVE AIRCRAFT CONTROL SYSTEMBACKGROUND OF THE INVENTION 1. Field of the Invention This inventionrelates to flight control systems and particularly to systems forcontrolling an aircraft over a wide range of flight conditions. Moreparticularly, this invention relates to an aircraft control systemhaving stability augmentation means including a constant high-gainfeedback loop for forcing response of the craft to follow a desiredresponse.

2. Description of the Prior Art Most present day high-performanceaircraft require stability augmentation systems tolerant of widevariations in flight conditions to bring the handling qualities of thecraft to desired specification levels. Prior to the present inventionthe systems included a computer for computing the high frequency gain ofthe craft commensurate with changing flight conditions and a variablegain element for changing control loop gain as a function of thecomputed information. Systems of this kind are termed "active adaptive"systems and are rather complex suffering decreased reliability with theapplication of redundancy techniques which are a necessary consequenceof present day fail-operative flight requirements.

SUMMARY OF THE INVENTION The device of the present invention is apassive adaptive" system in that it provides a constant high frequencygain to provide adequate stability over a wide range of flightconditions. A force sensor provides a signal corresponding to the forceapplied by the pilot to control the craft, and which signal is shaped bya model to provide an output corresponding to the desired response ofthe craft. The model output is combined with feedback signalscorresponding to the actual response of the aircraft, and which feedbacksignals have been blended in accordance with a predetermined gain. Thecombined signal is applied to a constant gain device and thegainadjusted combined signal is applied to a filter which provides ahigher gain in the frequency range of the desired response. The outputof the filter is applied to a damper servo for providing a correspondingmechanical output, and which mechanical output is combined with thepilot-applied inputs transmitted through a mechanical control system.The combined mechanical output drives a surface actuator which in turnoperates an aircraft control surface.

One object of this invention is to provide an aircraft control systemincluding stability augmentation means for controlling an aircraft overa wide range of flight conditions.

Another object of this invention is to provide stability augmentationmeans including a constant high-gain feedback loop for forcing theresponse of the craft to follow a desired response.

Another object of this invention is to provide a feedback loop includingaircraft acceleration sensors, and wherein said sensors are discretelypositioned forward of the center of gravity of the craft for sensingangular as well as linear acceleration to permit a higher feedback loopgain then would otherwise be the case.

Another object of this invention is to provide frequencyshaping meansfor permitting higher gains in the frequency range of the desiredaircraft response without compromising aircraft stability in otherfrequency ranges.

Another object of this invention is to blend various feedback gains topermit a sufficiently high constant gain for insuring that even atinsensitive flight conditions the aircraft response is forced to followthe desired response.

The foregoing and other objects and advantages of the invention willappear more fully hereinafter from a consideration of the detaileddescription which follows, taken together with the accompanying drawingswherein several embodiments of the invention are illustrated by way ofexample. It is to be expressly understood, however, that the drawingsare for illustration purposes only and are not to be construed asdefining the limits of the invention.

DESCRIPTION OF THE DRAWINGS FIG. I is a block diagram of a systemaccording to the invention for controlling an aircraft about its pitchaxis.

FIG. 2 is a block diagram of the device according to the invention forcontrolling an aircraft about its roll and yaw axes.

DESCRIPTION OF THE INVENTION FIG. I shows a force sensor 2 for sensing aforce F applied by a human or automatic pilot to control an aircraftabout its pitch axis and for providing a signal corresponding to thesensed force. Force F, is applied through suitable mechanical linkage 4to a mechanical summing means 6.

The signal provided by force sensor 2 is applied to a model 8. Model 8includes a simple second order filter network and provides a signalcorresponding to the desired dynamic response of aircraft 10 about itspitch axis. The signal from model 8 is applied to an electrical summingmeans 12.

A pitch rate gyro 14 mounted on aircraft I0 provides a signal 6corresponding to the pitch rate of the craft. The signal from gyro I4 isapplied to a gain device I6 having a predetermined constant gain K andwhich gain device 16 may be a simple fixed-gain amplifier and voltagedivider. The

signal from gain device 16 is applied to an electrical summing means I8.

A normal accelerometer 22 discretely mounted forward of the center ofgravity of aircraft l0 senses angular as well as linear acceleration andprovides a corresponding signal having a gain adjustment in accordancewith the location of accelerometer 22 relative to the center of gravityof the craft, and which signal is designated as I(.,6+a,,. The signalfrom accelerometer 22 is applied to summing means 18 and summed therebywith the signal from gain device 16.

The signal from summing means 18 is applied to summing means 12 andsummed thereby with the signal from model 8. The signal from summingmeans I2 is applied to a gain device 24 having a predetermined constantgain K, and which gain device 24 may be a simple fixed gain amplifierand voltage divider. The signal from gain device 24 is applied to alag-lead network 26 which shapes the signal for providing higher gain inthe frequency range of the desired aircraft response without increasinghe gain in other frequency ranges.

The signal from lag-lead network 26 is applied to a pitchdamper servo28. Pitch-damper servo 28 provides a corresponding mechanical outputwhich is applied to summing means 6 and summed thereby with themechanical input applied through linkage 4. Summing means 6 provides acombined mechanical output which is applied to a surface actuator 30 foroperating a pitch control surface such as the aircraft elevators (notshown) to control aircraft 10 about the pitch axis.

FIG. 2 shows a force sensor 31 for providing a signal corresponding to aforce F exerted by the human or automatic pilot to control aircraft 10about the roll axis. Force F is applied through suitable mechanicallinkage 46 to a mechanical summing means 44. The signal from forcesensor 31 is applied to a model 32. Model 32 includes a first-orderfilter network and provides a signal corresponding to the desireddynamic response of aircraft 10 about the roll axis. The signal frommodel 32 is applied to an electrical summing means 34.

A roll rate gyro 38 mounted on aircraft 10 provides a signal 0corresponding to the roll rate of the craft. The signal from gyro 38 isapplied to summing means 34 and summed thereby with the signal frommodel 32. The summed signal is applied to a gain device 36 having apredetermined constant gain K, and which gain device 36 may be a simplefixed-gain amplifier and voltage divider. The signal from gain device 36is applied to a lag-lead network 40 which shapes the signal forproviding high gain in the frequency range of the desired aircraftresponse without increasing gain in other frequency ranges.

The signal from lag-lead network 40 is applied to a rolldamper servo 42.Roll-damper servo 42 provides a corresponding mechanical output which isapplied to summing means 44 and summed thereby with the mechanicalinputs applied through linkage 46. Summing means 44 provides a combinedmechanical output which is applied to a surface actuator 48 foroperating a roll control surface such as the aircraft ailerons (notshown) to control aircraft 10 about the roll axis.

A lateral accelerometer 50 discretely mounted forward of the center ofgravity of aircraft l senses angular as well as linear acceleration andprovides a corresponding signal having a gain adjustment in accordancewith the location of accelerometer 50 relative to the center of gravityof the craft, and which signal is designated as K.-.t{j,tl1,,. Thesignal from accelerometer 50 is applied to a summing means 54.

A yaw-rate gyro 52 mounted on aircraft provides a signal It:corresponding to the yaw rate of the craft. The signal from gyro 52 isapplied to a gain device 53 having a predetermined constant gain K';,and which gain device 53 may be a simple fixed-gain amplifier andvoltage divider. The signal from gain device 53 is applied to summingmeans 54 and summed thereby with the signal from accelerometer 50. Thesignal from summing means 54 is applied to a gain device 56 having aconstant gain K", and which gain device 56 may be a simple fixed-gainamplifier and voltage divider.

The signal from gain device 56 is applied to a yaw-damper servo 58 whichprovides a corresponding mechanical output, and which output is appliedto a mechanical summing means 60 and combined thereby with mechanicalinputs applied through linkage 57 when the human or automatic pilotexerts a force F? to control aircraft 10 about its yaw axis. Thecombined output from summing means 60 is applied to a surface actuator64 for operating a yaw-control surface such as the aircraft rudder (notshown) to control aircraft 10 about the yaw axis.

OPERATION The system of the invention provides a high-gain feedback loopto force the response of aircraft l0 (acceleration) to follow a desiredresponse. Unlike the more complex "active adaptive systems whichcontinuously compute high frequency gain consistent with systemstability, the device of the present invention acts in a passive mannerand uses a constant high gain throughout the flight of the craft. Thus,the system does not require a complex computer and the variable-gainelement of the active system can be replaced by simple amplifiers andvoltage dividers. The level of the constant gain is set to providemaximum-feedback loop gain commensurate with stability margins at themost sensitive flight conditions.

in the case of a high-performance supersonic aircraft having a largerange of sensitivity, it would seem that a constant-feedback loop gainwhich insures adequate stability at the most sensitive flight conditionsdoes not provide a sufficiently tight feedback loop at insensitiveflight conditions. This is particularly true when the controlledvariable is acceleration as in the present case.

If, however, accelerometers 22 and S0 in FIGS. 1 and 2, respectively,are discretely positioned forward ofthe center of gravity of aircraftl0, angular as well as linear acceleration is sensed. The accelerationsignal takes on an anticipatory character to permit a much higherfeedback loop gain than would be possible if the accelerometers werelocated at the aircraft's center of gravity. Hence, even for highperformance aircraft, a sufficiently high constant gain can be employedso that at insensitive flight conditions the system performance can bemade to satisfy stringent handling specifications.

Lag-lead networks 26 and 40 in FIGS. 1 and 2, respectively, providehigher loop gains in the frequency range of the dynamic response ofaircraft 10 without increasing gain in other frequency ranges. Thus,system performance is improved without appreciably compromisingstability margins.

The use of lag-lead networks 26 and 40 as described above, together withjudicious blending of the various feedback gains, permits the employmentof a sufficiently high constant gain so that aircraft response followdesired response even at insensitive flight conditions.

Although but a single embodiment of the invention has been illustratedand described in detail, it is to be expressly understood that theinvention is not limited thereto. Various changes may also be made inthe design and arrangement of the parts without departing from thespirit and scope of the invention as the same will now be understood bythose skilled in the art.

I claim:

I. An aircraft control system comprising:

means for providing a signal corresponding to the force applied forcontrolling the craft about a flight axis;

means connected to the force signal means for shaping the signaltherefrom and for providing a signal corresponding to the desiredresponse of the craft;

means for providing feedback signals corresponding to the actualresponse of the craft;

means for combining the desired response signal and the feedbacksignals;

means for adjusting the gain of the combined signal by a predeterminedconstant factor;

means for filtering the gain-adjusted combined signal for increasing thegain in the frequency range of the desired response of the craft withoutincreasing the gain in other frequency ranges;

means responsive to the filtered gain-adjusted combined signal forproviding a mechanical output corresponding thereto;

means for providing a mechanical output corresponding to the forceapplied to control the craft about a flight axis; and

means responsive to the mechanical outputs for controlling the craft.

2. An aircraft control system as described by claim 1, wherein:

the first mentioned means provides a signal corresponding to the forceapplied to control the craft about the pitch axis;

the means for providing feedback signals corresponding to the actualresponse of the craft includes a pitch-rate gyro and a normalaccelerometer;

means are provided for adjusting the gain of the signal from thepitch-rate gyro; and

the normal accelerometer is discretely positioned forward of the centerof gravity of the craft for sensing angular as well as linearacceleration and for providing a corresponding signal having apredetermined gain adjustment in accordance with said position.

3. An aircraft control system as described by claim 2, wherein the meansresponsive to the mechanical outputs for controlling the craft includes:

means for combining said mechanical outputs; and

a surface actuator connected to the combining means and to aircraftelevators for actuating the elevators in response to the combinedmechanical outputs to control the craft about the pitch axis.

4. An aircraft control system as described by claim 2, including:

means for combining the gain-adjusted pitch rate signal and the normalacceleration signal corresponding to angular as well as linearacceleration and having a predetermined gain adjustment in accordancewith the discrete position of the accelerometer.

5. An aircraft control system described by claim 1, wherein:

the first mentioned means provides a signal corresponding to the forceapplied to control the craft about the roll axis; and

the means for providing feedback signals corresponding to the actualresponse of the craft includes a roll-rate gyro.

6. An aircraft control system as described by claim 5, wherein the-meansresponsive to the mechanical outputs for controlling the craft includes:

means for combining said mechanical outputs; and

a surface actuator connected to aircraft ailerons for actuating theailerons in response to the combined mechanical output.

7. An aircraft control system as described by claim 5, in-

cluding:

means for providing feedback signals corresponding to the actualresponse of the craft about the yaw axis;

means responsive to the feedback signals for providing a mechanicaloutput corresponding thereto;

means for providing a mechanical output corresponding to the forceapplied to control the craft about the yaw axis; and

means responsive to said mechanical outputs for controlling the craftabout the yaw axis.

8. An aircraft control system as described by claim 7,

wherein:

the means for providing feedback signals corresponding to the actualresponse of the craft about the yaw axis includes a yaw-rate gyro and alateral accelerometer;

means are provided for adjusting the gain of the signal from theyaw-rate gyro; and

the lateral accelerometer is discretely positioned forward of the centerof gravity of the craft for sensing angular as well as linearacceleration and for providing a cor responding signal having apredetermined gain adjustment in accordance with said position. 9. Anaircraft control system as described by claim 8, including:

means for combining the gain-adjusted yaw-rate signal and the lateralaccelerometer signal corresponding to angular as well as linearacceleration and gain adjusted in accordance with the location of theaccelerometer. 10. An aircraft control system as described by claim 9,including:

means for adjusting the gain of the combined signal. It. An aircraftcontrol system as described by claim 7, wherein:

the means responsive to said mechanical outputs for controlling thecraft about the yaw axis includes: means for combining said mechanicaloutputs; and a surface actuator connected to the aircraft rudder foractuating the rudder in response to the combined mechanical output.

